JP4150199B2 - gas turbine - Google Patents

Description

[0001]
BACKGROUND OF THE INVENTION
The present invention relates to a gas turbine including a large number of moving blades arranged on a turbine shaft and arranged in the shape of a moving blade row, and a plurality of stationary blades arranged in a turbine casing and arranged in the shape of a stationary blade row. .
[0002]
[Prior art]
Gas turbines are used in many fields to drive generators or work machines. In that case, the energy contained in the fuel is used to rotationally drive the turbine shaft. For this purpose, the fuel is burned in a combustion chamber to which compressed air compressed by an air compressor is supplied. The high-temperature and high-pressure working medium generated by the combustion of fuel in the combustion chamber is guided through a turbine device that is connected downstream of the combustion chamber, and expands while working there.
[0003]
In order to generate the rotational motion of the turbine shaft, a large number of blades arranged in the shape of blade groups or cascades are usually arranged on the turbine shaft. The moving blade receives the impact force from the working medium and drives the turbine shaft. In order to guide the working medium in the turbine device, a stationary blade row is usually arranged between adjacent moving blade rows, which are fixed to the turbine casing.
[0004]
When designing such a gas turbine, it is usually a design goal to obtain a particularly high efficiency in addition to the power produced. The increase in efficiency is basically achieved by increasing the outlet temperature of the combustion chamber for thermodynamic reasons. The working medium exits the combustion chamber at its outlet temperature and flows into the turbine device. Thus, for such gas turbines, it is desirable and achieved that the temperature of the working medium be between about 1200 ° C and 1300 ° C.
[0005]
However, when the working medium is at such a high temperature, the components and parts that are exposed to the working medium are subject to significant heat loads. Nevertheless, in order to guarantee a relatively long life of the component with high reliability, it is usually considered to cool the component (especially the moving blades and / or stationary blades of a turbine unit). . Therefore, turbine blades are generally designed to be coolable, and in particular, it must be ensured that the first cascade as viewed in the direction of flow of the working medium is effectively and reliably cooled. Each turbine blade typically has a coolant passage integrally formed with the blade or airfoil for cooling. The coolant is accurately guided from the coolant passage to a portion where the turbine blade is thermally heavily loaded.
[0006]
Generally, cooling air is used as the coolant. The cooling air is typically directed to each turbine blade in an open cooling manner through an integrally formed coolant passage. The cooling air flows out of the turbine blades and is mixed with the working medium guided in the turbine apparatus. However, the design output of such an open cooled gas turbine is limited to a low level. This is because, generally in connection with the limited mechanical load capacity of the individual components in the gas turbine, an increase in output can only be achieved by increasing the fuel supply. An increase in fuel supply per se means that the amount of coolant required to cool the turbine blades increases considerably, which also means a loss of useful compressor mass flow. The loss is only acceptable to a limited extent.
[0007]
[Problems to be solved by the invention]
An object of the present invention is to provide a gas turbine in which turbine blades are reliably cooled and suitable for a particularly high design output.
[0008]
[Means for Solving the Problems]
The problem is that, in the gas turbine of the type described at the beginning, the coolant passage is sufficient for the first partial passage and the second partial passage, and the first partial flow of the coolant flowing through the first partial passage is sufficient for the stationary blade. It is solved by being split so that the second partial flow of coolant that is used for cooling and flows through the second partial path is relayed almost without loss.
[0009]
The present invention starts from the idea that in gas turbines, particularly large design outputs can be obtained with high efficiency, since losses in useful compressor mass flow are kept particularly small. In order to keep the losses particularly small, gas turbines must be designed with a particularly small amount of coolant (especially cooling air). This is achieved by abandoning the open cooling scheme and utilizing a closed circuit cooling scheme. In this closed circuit cooling system, the coolant is introduced into the combustion process after passing through the turbine blades to be cooled. However, in such a closed circuit cooling system, a pressure loss occurs in the coolant circuit, and the pressure loss probably inhibits coolant flowing out of the turbine blades from being supplied to the combustion chamber of the gas turbine. It becomes so.
[0010]
On the one hand, in order to keep the pressure loss particularly small and thus possible for the first time anyway, in principle, a closed-circuit cooling system that is closed in the turbine blades, the coolant is divided inside the stationary blades and the coolant The first partial flow sufficiently cools the vane and a second partial flow of coolant is directed through the vane with substantially no loss. In this way, on the one hand the vane itself is supplied with sufficient coolant, and on the other hand another cooled vane is provided with sufficient coolant as well with almost no pressure loss. Dividing the coolant into two partial streams means parallel cooling of the original vane and the next turbine stage vane with respect to the coolant pressure. By abandoning the concept of continuously flowing coolant through the vanes of adjacent turbine stages and the guide wheels connecting these vanes, the coolant passage is guided to reduce the pressure loss in the coolant. It becomes possible to do. In addition, since the coolant passage is divided in the stationary vane, the guidance of the coolant is simplified. In addition, the flow path that is important for pressure loss is shortened, thus reducing the pressure loss in the coolant, thereby reducing the amount of coolant.
[0011]
In that case, a sufficient amount of coolant is suitably supplied to the airfoils of the respective stationary blades. In other words, the airfoil portions of the stationary blades forming the first stationary blade row are subjected to a particularly large heat load, and the amount of coolant necessary for sufficient cooling is particularly large here. Accordingly, the coolant is accurately supplied to this part, and the flow path for the coolant flow can also be made particularly short. Even when a relatively large amount of coolant is supplied to the airfoil, the pressure loss can be kept small, so that the coolant flowing out of the airfoil can be returned to the combustion process of the gas turbine.
[0012]
In an advantageous embodiment of the invention, the coolant guided in the second partial passage is used for cooling in the stationary blades of the second stationary blade row. For this purpose, the second partial passage is advantageously connected at the outlet side to a cooling passage system which guides the coolant to the stationary blades of the second stationary blade row without sufficient loss. The coolant flowing through the second partial path is fully utilized for cooling the stationary blades of the second stationary blade row. The “used” coolant for the stator vanes of the first and second vane rows is then directed to another coolant passage in the combustion chamber of the gas turbine. Thereby, the second partial flow of coolant only flows through the stationary blades of the second stator blade row, so that the pressure drop in the coolant is only small.
[0013]
Exactly, the necessary amount of coolant can be accurately supplied to the stationary blades forming the second stationary blade row through the second partial path. In particular, since the second stator blade row is exposed to a lower temperature than the first stator blade row, the amount of coolant used to cool the stator blades forming the second stator blade row is relatively small. Therefore, the pressure loss is kept small despite the relatively long flow path of the coolant in the second partial path.
[0014]
In an advantageous embodiment of the invention, the pre-connected is such that the pressure of the “used” coolant before returning to the flow chamber surrounding the combustion chamber is approximately the same as the pressure applied to the flow chamber. It is proposed to design a new coolant compressor. Such an arrangement ensures that the coolant utilized in the vane has a sufficiently large pressure level and that most of the required coolant is returned to the combustion process of the gas turbine. Indeed, by returning the necessary cooling air to the combustion process, a particularly high power increase and / or a large reduction in the amount of harmful substances generated is achieved.
[0015]
The advantages obtained by the invention are in particular the precise supply of coolant to the component to be cooled with a relatively small pressure loss by branching the flow path of the cooling air in the vanes forming the first vane row. There is to be able to do. In this way, the pressure loss that occurs in a limited range is compensated by a coolant compressor that is pre-connected, in particular on the coolant side. As a result, each stationary blade is cooled by a closed circuit cooling system, maintaining high efficiency, increasing output, reducing the amount of harmful substances generated, and reusing coolant (cooling air) in the combustion process be able to. The cooling air flow is divided as necessary in the stationary blades forming the first stationary blade row, and the first partial flow is used to cool the airfoil portion of the stationary blade with an appreciable pressure loss. A two-part flow passes through this vane with almost no pressure loss. As a result, the second partial flow is utilized to cool the stationary blades forming the second stationary blade row with a generally appreciable pressure loss. With regard to the pressure state of the coolant, it is closed by a general harmony between the coolant compressor connected in front and the coolant passages divided to parallelly cool the vanes of adjacent turbine stages. A circuit cooling scheme can be utilized to maintain the required cooling and reduce the amount of coolant required.
[0016]
DETAILED DESCRIPTION OF THE INVENTION
Hereinafter, embodiments of the present invention will be described in detail with reference to the drawings. In the drawings, the same parts are denoted by the same reference numerals.
[0017]
A gas turbine 1 in FIG. 1 includes a compressor 2 for combustion air, a combustion chamber 4, and a turbine 6 for driving the compressor 2 and a generator or a work machine (not shown). For this purpose, the turbine 6 and the compressor 2 are arranged on a common turbine shaft 8, also called a turbine rotor. The turbine shaft 8 is supported so as to be rotatable about its central axis 9 and is coupled to a generator or a work machine.
[0018]
The combustion chamber 4 is equipped with a number of burners 10 for burning liquid fuel or gaseous fuel. Further, the combustion chamber 4 is provided with a heat shielding element (not shown) on its inner wall.
[0019]
The turbine 6 has a number of blades 12 coupled to a turbine shaft 8. These blades 12 are annularly arranged on the turbine shaft 8, thereby forming a plurality of blade rows. The turbine 6 has a large number of stationary blades 14. These stationary blades 14 are similarly arranged in an annular shape, form a stationary blade row, and are fixed to an internal casing 16 of the turbine 6. The rotor blade 12 is used to drive the turbine shaft 8 under the impact force of the working medium M flowing through the turbine 6. On the other hand, the stationary blade 14 is used to guide the working medium M between two moving blade rows or moving blade rings that are continuous in the flow direction of the working medium M. One adjacent pair consisting of a ring (stator blade ring) or stator blade row of the stator blade 14 and a ring (moving blade ring) or rotor blade row of the rotor blade 12 is also called a turbine stage.
[0020]
Each stationary blade 14 has a blade base 18 also called a blade root portion. The blade base 18 is used to fix the stationary blade 14 to the inner casing 16 of the turbine 6 and is arranged as a wall element. The blade base 18 is a component that forms an outer boundary portion of the hot gas passage for the working medium M flowing through the turbine 6 and is heavily loaded thermally. Similarly, each rotor blade 12 is fixed to the turbine shaft 8 via a blade base 20 called a blade root portion.
[0021]
Guide wheels 21 are respectively disposed between the blade bases 18 of the stationary blades 14 of two stationary blade rows adjacent to each other with a space therebetween, and are fixed to the internal casing 16 of the turbine 6. Similarly, the inner surface of each guide wheel 21 is also exposed to a high temperature working medium M flowing through the turbine 6. The inner surface of the guide wheel 21 is spaced radially from the outer end 22 of the moving blade 12 positioned opposite thereto, and a gap exists between the two. The guide wheel 21 arranged between two adjacent stationary blade rows in particular protects the internal casing 16 or other casing inclusions from thermal overload by the hot working medium M flowing through the turbine 6. Used as an element.
[0022]
The gas turbine 1 is designed for a relatively high outlet temperature of about 1200 ° C. to 1300 ° C. of the working medium M flowing out of the combustion chamber 4 in order to obtain a relatively high efficiency. In order to enable this, at least some of the moving blades 12 and the stationary blades 14 are designed to be cooled by cooling air as a cooling medium. In order to clarify the flow path of the cooling air, the part immediately following the combustion chamber 4 of the gas turbine 1 is shown enlarged in FIG. As can be understood from this, the working medium M flowing out of the combustion chamber 4 first flows into a large number of stationary blades 14 forming a so-called first stationary blade row. These stationary blades 14 are suspended in the combustion chamber 4 via respective blade bases 18. When viewed in the flow direction of the working medium M, the stationary blades 14 that form the first stationary blade row, the stationary blades 14 that form the first stationary blade row, the stationary blades 14 that form the second stationary blade row, Followed by a rotor blade 12 forming a two rotor blade row.
[0023]
In order to enable a high design output in the gas turbine 1 having a particularly high efficiency, the gas turbine 1 is designed for a closed circuit cooling system of the stationary blades 14 forming the first stationary blade row and the second stationary blade row, respectively. Has been. In that case, it is considered that “used” cooling air flowing out from the above-described stationary blades 14 is returned to the combustion chamber 4 of the gas turbine 1. In order to suppress the pressure loss of the cooling air generated by such a cooling method sufficiently small and return the cooling air to the combustion chamber 4, the first stationary blade row is formed as viewed in the flow direction of the working medium M. Each stationary vane 14 has an integrally formed coolant passage, each symbolically indicated by an arrow 30. The inlet 32 of the coolant passage is arranged at the free end 34 of each stationary blade 14. Thus, when the coolant K is supplied to the coolant passage integrally formed with the stationary blades 14, the coolant flows through the respective stationary blades 14 from the free ends 34 toward the blade root portion, that is, the blade base 18.
[0024]
Inside the vane body of each stationary vane 14, the coolant passage is divided into a first partial path symbolically indicated by a first partial flow arrow 36 and a second part, as indicated by the division of the arrow 30. It branches off to the second partial path shown symbolically by the flow arrow 38. These partial streams are formed in such a way that coolant K can be precisely supplied with a relatively small pressure loss through them to particularly important points of the gas turbine 1. A first partial path symbolically indicated by a first partial flow arrow 36 is formed to cool the airfoil portion of each stationary blade 14. For this purpose, the stationary blade 14 has a cavity into which a partial flow of the coolant K is introduced, for example. The first partial passage is connected at its outlet side to a passage system communicating with a flow chamber 42 surrounding the combustion chamber 4, as symbolically indicated by the flow arrow 40. The flow chamber 42 is supplied with a mass flow rate of the compressor flowing out from the compressor 2 of the gas turbine 1. Therefore, when the partial flow of the coolant K guided through the first partial path is fed back to the flow chamber 42, the partial flow and the compressor mass flow are mixed. The mixed gas existing in the flow chamber 42 is subsequently supplied to the combustion process of the gas turbine 1.
[0025]
The second partial path symbolically indicated by the second partial flow arrow 38 inside the blade body of each stationary blade 14 is guided through the following guide wheel 21 on the outlet side when viewed in the flow direction of the working medium M. Connected to the cooling passage system. In this cooling passage system, the outlet side communicates with a coolant passage formed integrally with the stationary blade 14 that forms the second stationary blade row when viewed in the flow direction of the working medium M. This creates a flow path for the coolant K guided in the second partial path. This flow path first extends in the original vane in response to the second partial flow arrow 38, from there through the guide wheel 21, as indicated by the flow arrow 44, and the flow arrow 46. As shown by, the other stator blades 14 belonging to the second stator blade row meander and extend. From the outlet of the flow path, a partial flow of coolant K is similarly returned to the flow chamber 42 as indicated by flow arrows 48.
[0026]
By introducing the coolant K for the stationary blade 14 in this way, the coolant K that is particularly suited to the supply is supplied. The stationary blades 14 forming the first stationary blade row as viewed in the flow direction of the working medium M are subjected to a relatively large heat load. Accordingly, a relatively large amount of coolant K is supplied to the stationary blades 14 of the first stationary blade row. Since the flow path until returning to the flow chamber 42 is particularly shortened, the pressure loss can be kept small even when a relatively large amount of the coolant K is used. On the other hand, the coolant K is also supplied to the stationary blades 14 forming the second stationary blade row as viewed in the flow direction of the working medium M. In this case, the pressure loss is kept small here even though a relatively long flow path exists. In particular, this is because the partial flow of the coolant K is not involved in the cooling of the first stator blade row as viewed in the flow direction of the working medium M, and therefore there is almost no pressure loss in the stator blade 14 of the first stator blade row. Because it flows into As a result, the pressure loss of the coolant K as a whole is kept particularly small, so that the return of the coolant K to the flow chamber 42, and hence the gas turbine 1 is not started until the stationary blade 14 is cooled by the closed circuit cooling system. Return to the combustion process is possible.
[0027]
In order to particularly reliably maintain the pressure level required for such return of the coolant K, cooling is performed on the coolant side on the stator blades 14 forming the first stator blade row as viewed in the flow direction of the working medium M. A material compressor 50 is connected in front. In this embodiment, the coolant compressor 50 is formed as an axial compressor, with a number of compressor elements 52 rigidly coupled to the turbine shaft 8 and a number of compressors rigidly coupled to the turbine casing 6. Element 54. The coolant compressor 50 is supplied with coolant K on the inlet side, and the outlet side is provided with the coolant K of the stationary blade 14 via the coolant passage 58 defined by the turbine shaft 8 and the passenger compartment compressor element 56. Communicated with the inlet 32. In order to supply the coolant K, the coolant compressor 50 is connected to the compressor 2 on the inlet side. As a result, a part of the compressor mass flow rate flowing out of the compressor 2 is branched as the coolant K for the stationary blade 14.
[0028]
The coolant compressor 50 is designed in accordance with a predetermined operation state in which the coolant K for the stationary blade 14 is returned to the combustion process of the gas turbine 1. Therefore, the coolant compressor 50 reduces the pressure applied to the flow chamber 32 when the coolant K returns to the flow chamber 42 after flowing through the first partial path or the second partial path and the subsequent flow path. It is designed to increase the pressure of the coolant K at its designed operating conditions so that it has approximately.
[Brief description of the drawings]
FIG. 1 is a schematic longitudinal sectional view of a gas turbine.
2 is a schematic explanatory diagram of a cooling passage system according to the present invention in the gas turbine of FIG. 1. FIG.
[Explanation of symbols]
1 Gas turbine 8 Turbine shaft 12 Rotor blade 14 Stator blade

Claims (3)

In a gas turbine that includes a turbine casing and a stator blade row, a coolant passage is integrally formed with the stator blades of the stator blade row, and a cooling circuit through which the coolant flows is connected to the coolant passage. In the first partial path and the second partial path, the first partial flow of the coolant flowing through the first partial path is used to sufficiently cool the stationary blade, and the second of the coolant flowing through the second partial path. The coolant is divided so that the partial flow is relayed almost without loss, and the first stationary blade row is divided into the first partial passage and the second partial passage when viewed in the flow direction of the working medium. A gas turbine comprising a stationary blade having a passage . The second partial stream of coolant, almost without losses, as seen in the direction of flow of the working medium, characterized in that it is Tsugioku to vanes of the next stator blade row according to claim 1 Symbol placement of the gas turbine. The coolant compressor pre-connected to the cooling circuit increases the pressure of the coolant so that it almost has the pressure applied to the flow chamber as it returns to the flow chamber surrounding the combustion chamber of the gas turbine. according to claim 1 or 2, wherein the gas turbine, characterized in that it is designed to.

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